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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 631 records · Page 35
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Laser power beaming for rocket propulsion and airbreathing propulsion

The developing technology of laser power beaming is introduced, and two systems are used as examples of the capabilities of the laser for beamed energy. In the first system, the potential of the laser to power flight systems ranging from hypersonic air-breathing launch vehicles to commercial jet transports is examined. Attention is given to the possibility of an air-breathing propulsion which offers the promise of a global air transportation network independent of kerosene and powered by solar energy. In addition, consideration is given to a new type of rocket propulsion based on the laser's ability to concentrate coherent laser energy to high power densities. Focused laser beams would heat the propellants directly to produce specific impulses approaching ion and MHD rocket levels, and would do so without the burden of a heavy electrical power supply.

Hertzberg, A.↗
NASA NTRS1981

Primary propulsion/large space system interaction study

An interaction study was conducted between propulsion systems and large space structures to determine the effect of low thrust primary propulsion system characteristics on the mass, area, and orbit transfer characteristics of large space systems (LSS). The LSS which were considered would be deployed from the space shuttle orbiter bay in low Earth orbit, then transferred to geosynchronous equatorial orbit by their own propulsion systems. The types of structures studied were the expandable box truss, hoop and column, and wrap radial rib each with various surface mesh densities. The impact of the acceleration forces on system sizing was determined and the effects of single point, multipoint, and transient thrust applications were examined. Orbit transfer strategies were analyzed to determine the required velocity increment, burn time, trip time, and payload capability over a range of final acceleration levels. Variables considered were number of perigee burns, delivered specific impulse, and constant thrust and constant acceleration modes of propulsion. Propulsion stages were sized for four propellant combinations; oxygen/hydrogen, oxygen/methane, oxygen/kerosene, and nitrogen tetroxide/monomethylhydrazine, for pump fed and pressure fed engine systems. Two types of tankage configurations were evaluated, minimum length to maximize available payload volume and maximum performance to maximize available payload mass.

Coyner, J. V.↗
NASA NTRS1981

Effect of thrust chamber configuration on MPD arcjet performance

The performance of quasisteady multimegawatt MPD thrusters is significantly affected by anode thickness, location, and orifice radius, and by cathode length. Terminal voltage oscillations and electrode erosion are deferred until higher currents by anodes at more downstream locations and of smaller orifice radius and by cathodes of greater length. Without an optimized geometry, specific impulses of 3300 s and thrust efficiencies up to 31% are implied by the best data.

King, D. Q.↗
NASA NTRS1982

Mass driver reaction engine characteristics and performance in earth orbital transfer missions

Configurations of a typical mass driver reaction engine (MDRE) are presented and its use for delivery of payloads to geosynchronous orbit (GEO) from low earth orbit (LEO) is discussed. Basic rocket equations are developed for LEO to GEO round-trip missions using a single exhaust velocity. It is shown that exhaust velocities in the 5-10 km/sec range (specific impulse of 500-1000 sec) are well suited for mass drivers, minimizing the overall cost of missions. Payload delivery rate fractions show that there is little to be gained by stretching out LEO to GEO transfer times from 90 to 180 days. It therefore pays to use the shorter trip time, approximately doubling the amount of delivered payload during any fixed time of use of the MDRE.

Snow, W. R.↗
NASA NTRS1982

Advanced rocket propulsion technology assessment for future space transportation

Single-stage and two-stage launch vehicles were evaluated for various levels of propulsion technology and payloads. The evaluation included tradeoffs between ascent flight performance and vehicle sizing that were driven by engine mass, specific impulse, and propellant requirements. Numerous mission, flight, and vehicle-related requirements and constraints were satisfied in the design process. The results showed that advanced technology had a large effect on reducing both single- and two-stage vehicle size. High-pressure hydrocarbon-fueled engines that were burned in parallel with two-position nozzle hydrogen-fueled engines reduced dry mass by 23% for the two-stage vehicle and 28% for the single-stage vehicle as compared to an all-hydrogen-fueled system. The dual-expander engine reduced single-stage vehicle dry mass by 41%. Using advanced technology, the single-stage vehicle became comparable in size and sensitivity to that of the two-stage vehicle for small payloads.

Wilhite, A. W.↗
NASA NTRS1980

Analysis and design of ion thruster for large space systems

Design analyses showed that an ion thruster of approximately 50 cm in diameter will be required to produce a thrust of 0.5 N using xenon or argon as propellants, and operating the thruster at a specific impulse of 3530 sec or 6076 sec respectively. A multipole magnetic confinement discharge chamber was specified.

Poeschel, R. L.↗
NASA NTRS1982

Operation of the J-series thruster using inert gas

Electron bombardment ion thrusters using inert gases are candidates for large space systems. The J-Series 30 cm diameter thruster, designed for operation up to 3 k-W with mercury, is at a state of technology readiness. The characteristics of operation with xenon, krypton, and argon propellants in a J-Series thruster with that obtained with mercury are compared. The performance of the discharge chamber, ion optics, and neutralizer and the overall efficiency as functions of input power and specific impulse and thruster lifetime were evaluated. As expected, the discharge chamber performance with inert gases decreased with decreasing atomic mass. Aspects of the J-Series thruster design which would require modification to provide operation at high power with insert gases were identified.

Rawlin, V. K.↗
NASA NTRS1982

LOX/hydrocarbon auxiliary propulsion system study

Liquid oxygen (LOX)/hydrocarbon propulsion concepts for a "second generation' orbiter auxiliary propulsion system was evaluated. The most attractive fuel and system design approach identified, and the technology advancements that are needed to provide high confidence for a subsequent system development were determined. The fuel candidates were ethanol, methane, propane, and ammonia. Even though ammonia is not a hydrocarbon, it was included for evaluation because it is clean burning and has a good technology base. The major system design options were pump versus pressure feed, cryogenic versus ambient temperature RCS propellant feed, and the degree of OMS-RCS integration. Ethanol was determined to be the best fuel candidate. It is an earth-storable fuel with a vapor pressure slightly higher than monomethyl hydrazine. A pump-fed OMS was recommended because of its high specific impulse, enabling greater velocity change and greater payload capability than a pressure fed system.

Orton, G. F.↗
NASA NTRS1983

Evaluation of the use of on-board spacecraft energy storage for electric propulsion missions

On-board spacecraft energy storage represents an under utilized resource for some types of missions that also benefit from using relatively high specific impulse capability of electric propulsion. This resource can provide an appreciable fraction of the power required for operating the electric propulsion subsystem in some missions. The most probable mission requirement for utilization of this energy is that of geostationary satellites which have secondary batteries for operating at high power levels during eclipse. The study summarized in this report selected four examples of missions that could benefit from use of electric propulsion and on-board energy storage. Engineering analyses were performed to evaluate the mass saved and economic benefit expected when electric propulsion and on-board batteries perform some propulsion maneuvers that would conventionally be provided by chemical propulsion. For a given payload mass in geosynchronous orbit, use of electric propulsion in this manner typically provides a 10% reduction in spacecraft mass.

Poeschel, R. L.↗
NASA NTRS1982

Electric thruster research

It has been customary to assume that ions flow nearly equally in all directions from the ion production region within an electron-bombardment discharge chamber. In general, the electron current through a magnetic field can alter the electron density, and hence the ion density, in such a way that ions tend to be directed away from the region bounded by the magnetic field. When this mechanism is understood, it becomes evident that many past discharge chamber designs have operated with a preferentially directed flow of ions. Thermal losses were calculated for an oxide-free hollow cathode. At low electron emissions, the total of the radiation and conduction losses agreed with the total discharge power. At higher emissions, though, the plasma collisions external to the cathode constituted an increasingly greater fraction of the discharge power. Experimental performance of a Hall-current thruster was adversely affected by nonuniformities in the magnetic field, produced by the cathode heating current. The technology of closed-drift thrusters was reviewed. The experimental electron diffusion in the acceleration channel was found to be within about a factor of 3 of the Bohm value for the better thruster designs at most operating conditions. Thruster efficiencies of about 0.5 appear practical for the 1000 to 2000 s range of specific impulse. Lifetime information is limited, but values of several thousands of hours should be possible with anode layer thrusters operated or = to 2000 s.

Kaufman, H. R.↗
NASA NTRS1983

Electrothermal thruster diagnostics. Volume 1: Executive summary

A flight-qualified electrothermal thruster demonstrated its adaptability to a variety of propellants. Originally qualified for operation with hydrazine propellant, it was operated with nitrogen, hydrogen, and ammonia propellants, demonstrating 73, 61, and 52 percent overall efficiency with these propellants, respectively, when tested over a wide range of operating conditions. By introducing a preheater to admit hot, rather than cold, propellant inlet gases to the thruster's augmentation heat exchanger, delivered specific impulse closer to theoretical performance limits should be achieved.

Zafran, S.↗
NASA NTRS1983

Advanced LOX/H sub 2 engine technologies for future OTVs

The optimization methods and trade studies which were used to define a 1980 state-of-the-art design of orbital transfer vehicles are summarized. The advanced thrust chamber, turbomachinery, and engine power cycle technologies recommended for development and verification during the 1980's are discussed. The state-of-the-art design serves as one of the reference points for evaluation of the advanced technology concepts. A second reference point is provided by the 1960 state-of-the-art RL-10 engine. Advanced heat energy extraction concepts in the combustor and injector are presented that will permit the high chamber pressures expected of the advanced expander cycle engine concept. Advanced turbomachinery concepts are selected that will help utilize efficiently the heat energy extracted in the thrust chamber and aid in uprating chamber pressures to values five times those realized with 1960 to 1980 technologies. The higher chamber pressure benefits are maximized with nozzle concepts that will package within a short retracted length and extend to the very high expansion area ratios required for specific impulse values significantly higher than those now realized by operational systems. Control system component concepts studied will allow utilization of the high turbine drive energy levels in a reliable repeatable manner that will reduce program risk and enhance overall space transportation system mission life cycle cost.

Glass, J.↗
NASA NTRS1982

Magnetoplasmadynamic thruster development

Current research on self-field MPD thrusters presents the possibility of developing high power/low cost electric propulsion. The interest generated in this propulsion concept within both NASA and the U.S. Air Force has led to a coordinated interagency effort, and the recent completion of several test sites are expected to speed thruster concept development. Efficiencies approaching 40% at 2200 sec have been experimentally demonstrated. Among the characteristics recommending development of MPD thrusters are high thrust density and power handling capability, simplicity, and a wide dynamic range of thrust and specific impulse. MPD thrusters can, moreover, operate either continuously, at high available power levels, or in a pulsed mode, to match lower power levels.

Pawlik, E. V.↗
NASA NTRS1983

A reexamination of the laser supported combustion wave

Kantrowitz (1972) and Minovitch (1972) have proposed the use of laser sustained plasmas as a means to heat a rocket propellant. Recent studies of laser-powered propulsion have been directed toward the application of high-specific-impulse space propulsion systems for orbital transfer missions. Analyses of rocket performance relied heavily on the concept of the laser-supported combustion (LSC) wave. Raizer (1971) first drew the analogy between laser-sustained plasmas and combustion waves in an analysis. The Raizer model was later applied to hydrogen by Kemp and Root (1979). In connection with certain problems arising with the approach considered by Kemp and Root, the present investigation is concerned with a reexamination of the Raizer model. Attention is given to a numerical approach for the entire LSC wave in hydrogen, taking into account the incorporation of the proper boundary conditions far downstream of the wave.

Keefer, D.↗
NASA NTRS1983

Effects of tripropellant engines on earth-to-orbit vehicles

The effects of tripropellant engines on earth-to-orbit vehicles is examined in terms of their impact on engine configurations and launch capabilities. Hydrocarbon fuel with some oxygen is used in tripropellant fuels, with hydrogen as a back-up fluid for coolant and driving the pumps. Engine concepts which implement tripropellant fuels include a hydrogen-gas generator engine with staged combustion, a two-mode engine burning hydrocarbon fuel in one chamber and hydrogen in another, a dual expander engine with a central hydrocarbon nozzle and an annular hydrogen nozzle, and a dual throat engine. The use of hydrocarbons reduces the fuel weight by providing a higher specific impulse than with LOX-LH2 systems alone. Studies have shown that single-stage-to-orbit vehicles capable of lifting 13.6 Mg are possible with tripropellant engines. A dual-expander engine, is identified as offering the most fuel dry mass reduction for a given payload if cooling requirements can be satisfied. Further development of the tripropellant engines is concluded to be beneficial.

Martin, J. A.↗
NASA NTRS1983

A parametric study of thermally augmented O2/H2 rocket enines

An analysis is presented of the performance of a conceptual propulsion system in which liquid hydrogen and liquid oxygen are first vaporized and heated by thermal power from a spacecraft nuclear electric power supply prior to combustion in a rocket engine. Calculations of the specific impulse (I-sp) are presented for a series of O2/H2 oxidizer-to-fuel mixture ratios and reactant pre-heat temperatures (T-aug). It is found that the ratio of the augmentation power to the total engine jet power (P-aug/P-tot) determines the engine thrust for a given P-aug, T-aug, and I-sp. In addition, the performance of an unaugmented O2/H2 and a heated-hydrogen rocket engine were also calculated for the same conditions. Results show that an augmented O2/H2 engine has three to eight times the thrust of a heated H2 engine for a given P-aug. It is concluded that it should be possible to design a high/low thrust propulsion system using a common propellant, with the option of using thermal power from the nuclear electric power supply to augment the low-thrust propulsion system during cruise periods when the reactor's full electric power is not needed.

Frisbee, R. H.↗
NASA NTRS1983

Experimental Investigation of a Hall-Current Accelerator

The Hall-current accelerator is being investigated for use in the 1000-2000 sec. range of specific impulse. Three models of this thruster were tested. The first two models had three permanent magnets to supply the magnetic field and the third model had six magnets to supply the field. The third model thus had approximately twice the magnetic field of the first two. The first and second models differ only in the shape of the magnetic field. All other factors remained the same for the three models except for the anode-cathode distance, which was changed to allow for the three thrusters to have the same magnetic field integral between the anode and the cathode. These Hall thrusters were tested to determine the plasma properties, the beam characteristics, and the thruster characteristics. The thruster operated in three modes: (1) main cathode only, (2) main cathode with neutralizer cathode, and (3) neutralizer cathode only. The plasma properties were measured along an axial line, 1 mm inside the cathode radius, at a distance of 0.2 to 6.2 cm from the anode. Results show that the current used to heat the cathode produced nonuniformities in the magnetic field, hence also in the plasma properties. In a Hall thruster this general design appears to provide the most thrust when operated at a magnetic field less than the maximum value studied.

Plank, G. M.↗
NASA NTRS1984

Low loss injector for Space Shuttle main engine. Center director's discretionary fund

An efficient propellant injection method to raise the Space Shuttle Main Engine (SSME) thrust and payload is discussed. Relatively large diameter injector elements with low pressure loss are recommended for the main combustion chamber and the pre-burners. Smaller losses admit more propellant flow which then raises thrust. Payload is not only gained by specific impulse but also by thrust. The chamber pressure is stabilized by selecting the proper cavity size for the injector elements while reducing the injection pressure loss which normally is kept high for stability. The rather large injector element recesses provide acoustic damping which makes baffles and acoustic absorbers unnecessary. A tenfold reduction of flow induced stresses which are rather high in the present design is shown. Relaxed tolerances, fewer elements, and better maintenance are offered. The study was conducted under a center director discretionary fund assignment.

Vonpragenau, G. L.↗